Method and apparatus for measuring the structural integrity of a safe-life aircraft component

ABSTRACT

The structural integrity of a safe-life aircraft component on an aircraft is measured and assessed by a processing unit. The component includes a load-bearing metal element that is free from cracks. In the method, acoustic emissions generated in the metal element are converted into electronic signals. The acoustic emissions converted include relevant acoustic emissions resulting from changes in the structure of the element that make the element more susceptible to the formation of cracks. The electronic signals are set to a processing unit. The processing unit processes over time the signals in conjunction with stored reference data that allows a measure of the structural integrity to be made. Information providing a measure of the structural integrity of the aircraft component is outputted. Thus, deterioration of the structure of the component can be detected and monitored before a crack occurs.

RELATED APPLICATIONS

The present application is a continuation of U.S. application Ser. No.11/428,976, filed on Jul. 6, 2006 and is based on, and claims priorityfrom, UK Patent Application Number 0513901.9, filed Jul. 6, 2005, thedisclosure of which is hereby incorporated by reference herein in itsentirety.

BACKGROUND

The present invention relates to measuring or assessing the structuralintegrity of, the structural change in, or damage to, an aircraftcomponent, and in particular to a load bearing metal element used as asafe-life component on an aircraft.

A safe-life component on an aircraft must be structurally safethroughout the entire working life of the component and in particularmust not be allowed to develop cracks that could prejudice suchstructural safety, to yield to any non-negligible degree or to fail inany other way. (It will be understood that superficial micro-cracking onthe surface of a safe-life component does not in itself constitutecracking that could prejudice structural safety. As such in the contextof the present invention it will be understood that a component may beconsidered as being “crack-free” or “free from cracks”, despite thecomponent having superficial micro-cracking, provided that suchmicro-cracking is of a nature that does not in itself prejudicestructural safety.)

Safe-life components can be distinguished from “fail-safe” components,which are designed to be able to sustain damage or be allowed to developcracks, to yield, or even fail, without presenting an unacceptableshort-term safety risk. Fail-safe components may for example be able tosustain damage or partial failure without significantly affecting theability of the component to perform its function, or if so significantlyaffected there are other components that are able to act as a back-up.For example, the failure of a safe-life component during service wouldbe at risk of prejudicing safety to an unacceptable degree even in theshort-term, whereas failure of a fail-safe component would be able to betolerated until the next available opportunity for maintenance.Operators of aircraft typically have to conduct scheduled inspections ofsafe-life components at fixed intervals. Also, a safe-life componentgenerally has to be replaced after a certain length of service so as tomanage effectively the risk of failure of such a component in service.In view of the unacceptability of failure of a safe-life componentduring service, safe-life components are typically withdrawn far inadvance of the possible maximum length of useable service andconsequently the maximum length of service for safe-life components istypically conservatively short. Whilst such short lifetimes of safe-lifecomponents is wasteful, there is currently no means of effectivelyreducing the risk that a particular safe-life aircraft component willfail significantly in advance of the average lifetime. The presentinvention has been made in recognition that there has been a lack ofeffective means of assessing the structural integrity of such safe-lifecomponents once in use.

Various methods exist for measuring the structural integrity of a loadbearing metal element including, for example, non-destructive testingmethods (such as by means of X-ray radiography) and microscope analysis.Such techniques with their limited accuracy and other disadvantages,have limited application in relation to assessing and monitoring thestructural integrity of a safe-life component. The present inventionconcerns the use of acoustic emission monitoring to assess thestructural integrity of a safe-life aircraft component.

A method for detecting and monitoring fractures in a structure bymonitoring acoustic emissions is disclosed in International PatentApplication No. PCT/GB01/02213 (published under No. WO 01/94934) whichis incorporated herein by reference in its entirety. The method ofdetecting and monitoring for damage in metal structures as described inWO 01/94934 relies on monitoring acoustic emissions caused by fracturesor cracks. Such a method is therefore of no use when monitoring fordamage in safe-life aircraft components, because it is a requirementthat such components are free from cracks.

An experiment concerning the use of acoustic emissions to check thecondition of a cylindrical specimen at high temperature is described ina paper entitled “Acoustic-Emission Study of Damage Accumulation DuringAlternating Load-Cycle Loading at Elevated Temperature” by N. G. Bychkovet al which is translated from “Problemy Prochnosti, No. 11, pp. 21-23,November 1983” of the “Central Institute of Aircraft Engine Design,Moscow and which is incorporated herein by reference in its entirety.The teaching of that paper relates primarily to monitoring of crackgrowth, and predicting imminent fracture in high temperature samples, inan experimental/laboratory setting utilizing a waveguide to carryacoustic emissions from the sample, which is housed in a furnace, to anacoustic emission transducer. As mentioned above, the present inventionis concerned with structural health monitoring of safe-life componentson an aircraft, such components being required to be free from cracks.

SUMMARY

It is an aim of the present invention to provide an improved method ofmeasuring the structural integrity of a load bearing aircraft componentfor example by providing a method of measuring the structural integrityof a safe-life aircraft component made from a metal element.

The present invention provides a method of measuring the structuralintegrity of a safe-life aircraft component comprising a load-bearingmetal element, the method including the steps of:

converting acoustic emissions generated in the metal element intoelectronic signals, the acoustic emissions converted including “relevantacoustic emissions”, namely acoustic emissions resulting from changes inthe structure of the element that make the element more susceptible tothe formation of cracks,

sending the electronic signals to a processing unit, and,

with the processing unit, processing over time the signals inconjunction with stored reference data that allows a measure of thestructural integrity to be made from the signals sent to the processingunit, and

outputting information providing a measure of the structural integrityof the aircraft component. Thus the method of the invention may beimplemented to monitor for damage to, changes in, or deterioration of,the structure of the metal element before a crack occurs.

It will of course be appreciated that the electronic signals may bemodified before being received by the processor. For example, theelectronic signals may be converted from analogue to digital signals.

The step of processing the signals may include assessing whether one ormore signals correspond to a “relevant acoustic emission” and preferablyincludes assessing which of each acoustic emission signal corresponds toa “relevant acoustic emission”. For example, acoustic emissions thatconform to pre-set criteria characterising a relevant acoustic emissionmay be deemed as corresponding to a relevant acoustic emission. Acousticemission signals assessed as not corresponding to a relevant acousticemission are preferably discarded by the processing unit, for examplebeing ignored for the purposes of performing the method of the inventionas defined herein. The step of assessing whether an acoustic emissionmay be deemed as corresponding to a relevant acoustic emission mayinclude assessing whether the acoustic emission is one that is typicalof an acoustic emission resulting from changes in the structure of theelement at a scale of the order of a few microns or less.

The step of processing the signals advantageously includes calculating,and preferably additionally monitoring, the cumulative number ofrelevant acoustic emissions over a period, for example a pre-set period.The period may for example be a period of time measured from the firsttime the method is performed on the metal element. The period of timemay exclude times during which the metal element is not being subjectedto loading of the type likely to change or affect the structuralintegrity of the aircraft component. The period may alternatively bemeasured as a number of nominal cycles of fatigue loading. The periodmay be measured as a number of actual cycles of fatigue loading (thecycles being detected by an appropriately arranged sensing system). Sucha calculation may be performed without assessing the location of theacoustic emissions. In contrast to methods of the prior art (such asthat described in WO 01/94934), the present invention is able to measurethe global structural properties of an element without first having todetermine the locations of the source of the acoustic emissions in theelement. It will of course be appreciated that location information mayadditionally be used in the performance of a method according to thepresent invention.

The step of processing the signals may include a step of effectivelycomparing the cumulative number of relevant acoustic emissions with apre-set threshold. The threshold may for example be set such that abovethe threshold there is a given probability of a crack having occurred inthe element. Once the threshold is exceeded the method may includetaking further action, for example, remedial action.

The step of processing the signals may include additionally oralternatively include a step of calculating, and preferably additionallymonitoring, the number of relevant acoustic emissions over a pre-setperiod. Again, the pre-set period may be a pre-set period of time oralternatively a pre-set period of nominal loading cycles. The pre-setperiod is preferably short enough to be short relative to the expectedlifetime of the element, but long enough that the number of relevantacoustic emissions detected during said period can be considered as notbeing significantly affected by the inherently noisy and seeminglyrandom nature of the rate of emission of relevant acoustic emissions.The pre-set period may be a constant period of time in the past asmeasured from the instant at which the step of processing the signals isperformed. Again, the number of relevant acoustic emissions over theperiod may be compared to a threshold and if the threshold is exceededappropriate further action may be taken.

The step of processing the signals may include performing calculationsusing a measure of the size of a relevant acoustic emission, forexample, the peak amplitude, the average signal level, the rise time ofthe emission, the energy and/or the duration of the emission. Themeasure of the size of a relevant acoustic emission may be used toweight the relevance of the relevant acoustic emission as compared toother relevant acoustic emissions of a different size.

The method may include a step of detecting acoustic emissions atdifferent frequencies and/or calculating a characteristic relatingthereto. The step of processing the signals may for example includetaking account of the frequencies of the respective relevant acousticemissions. The frequency of the signal used in such a step may forexample be the fundamental frequency, or a frequency equal to an integermultiple of the fundamental frequency.

The step of processing the signals may include a step of calculating acharacteristic relating to (for example being equal to) the rate ofrelevant acoustic emissions.

The step of processing the signals may include a step of calculating acharacteristic relating to (for example being equal to) the spatialdensity of relevant acoustic emissions.

The step of processing the signals may include a step of calculating acharacteristic relating to (for example being equal to) an indication ofthe timing or order of successive relevant acoustic emissions.

The step of processing the signals may include weighting any of theparameters or characteristics mentioned herein according to the size ofthe relevant acoustic emissions.

The step of processing signals may include calculating one or moreproximity characteristics that provide an indication of the proximity ofthe sources of respective relevant acoustic emissions relative to eachother. Such proximity characteristics may simply be in the form oflocation co-ordinates. Such proximity characteristics may be associated,or embedded with, any of the other parameters or characteristicsmentioned herein. Thus, the step of processing the signals may include astep of taking into account an indication of the proximity of thesources of the respective relevant acoustic emissions relevant to eachother.

The step of processing the signals may include a step of differentiatinga first variable with respect to a second variable. For example, thedifferential calculated may be a measure of the rate of change of thecumulative number of relevant acoustic emissions.

The step of processing the signals may include a step of integrating afirst variable over a range defined by means of a second variable. Forexample, the step of processing the signals may include a step ofcalculating a characteristic relating to (for example being equal to)the integral of the cumulative number of relevant acoustic emissionswith respect to a pre-set period.

The step of processing the signals may include calculatingcharacteristics of notional graphs (such characteristics for exampleincluding differentials or integrals), the notional graphs having atleast two axes dependent on any two independent variables representativeof any of the parameters or characteristics mentioned herein. It will beunderstood of course that the characteristics of the notional graphs maybe calculated either with or without creating a physical representation(such as a drawing for example) of the graph.

The step of processing the signals may include using non-acoustic data.Such non-acoustic data may include a measure of stress and/or strain inrelation to the metal element as measured by for example one or morestrain gauges or other sensors associated with the metal element. Thenon-acoustic data may alternatively or additionally include a measure oftemperature. The method may of course combine any of the above-describedsteps of processing the signals. For example, both characteristicsrelating to the cumulative number of relevant acoustic emissions, and tothe number of relevant acoustic emissions over a pre-set period of timemay be calculated and monitored. In such a case, the method may monitorwhether the two characteristics meet given criteria, such as eachcharacteristic exceeding a respective given threshold. In the event thatsuch criteria are met the method may include taking further action. Moregenerally, the step of processing the signals may include a step ofcalculating a plurality of characteristics selected from any of thecharacteristics or parameters mentioned herein. The plurality ofcharacteristics so calculated may then be compared with pre-setcriteria, the criteria being selected such that once met the aircraftcomponent is deemed to be in need of further, for example remedial,action. The step of calculating a plurality of different characteristicsmay for example comprise performing two or more of steps (a) to (n),where steps (a) to (n) are as follows:

(a) assessing whether one or more signals correspond to a “relevantacoustic emission”,

(b) assessing which of each acoustic emission signal corresponds to a“relevant acoustic emission”,

(c) calculating the cumulative number of relevant acoustic emissionsover a period,

(d) assessing the location of the acoustic emissions,

(e) effectively comparing the cumulative number of relevant acousticemissions with a pre-set threshold,

(f) calculating the number of relevant acoustic emissions over a pre-setperiod,

(g) using a measure of the size of a relevant acoustic emission,

(h) detecting acoustic emissions at different frequencies,

(i) calculating a characteristic relating to the rate of relevantacoustic emissions,

(j) calculating a characteristic relating to the spatial density ofrelevant acoustic emissions,

(k) calculating a characteristic relating to the timing or order ofsuccessive relevant acoustic emissions,

(l) calculating one or more proximity characteristics that provide anindication of the proximity of the sources of respective relevantacoustic emissions relative to each other,

(m) calculating a characteristic relating to the integral of thecumulative number of relevant acoustic emissions with respect to apre-set period, and

(n) calculating characteristics of notional graphs, the notional graphshaving at least two axes dependent on any two independent variablesrepresentative of any of the parameters or characteristics mentioned insteps (a) to (m) above.

The method may of course also include implementing the further action.In such a case the step of implementing of the further action may ofcourse replace the step of deeming of the need for the further action.

Above it is stated that further action may be taken, or may be deemed tobe in need of being taken, if a threshold is exceeded or if certaincriteria are met. The further action may simply be to provide anindication, electronic or otherwise, that indicates that the thresholdhas been exceeded or the criteria have been met, as is appropriate. Suchan indication may be arranged to be arranged to be discoverable duringmaintenance procedures relating to the aircraft component. The furtheraction may comprise testing the aircraft component or a part thereof,for example the metal element with other means or performing such tests,if already routinely performed, more often. The further action mayinclude further non-destructive testing to obtain an assessment of thestructural integrity of the metal element. Preferably, suchnon-destructive testing is not dependent on the monitoring of acousticemissions.

The further actions may be in the form of physical activity affectingthe use or function of the aircraft component. For example, the furtheraction may comprise replacing the aircraft component, or a part thereof.

The further action may include a step of repairing the aircraftcomponent or a part thereof.

The criteria mentioned above may be pre-selected by means of a methodincluding performing empirical testing, for example by means of a seriesof experiments using test rigs. The criteria may additionally oralternatively be pre-selected by means of a method including performingcomputer modelling of the aircraft component.

It will be appreciated that the thresholds and criteria mentioned hereinmay form the totality or part of the stored reference data that is usedwhen processing the signals to provide the information outputted by theprocessing unit that provides the measure of the structural integrity ofthe element.

The information outputted may include information in any of a widevariety of forms. The outputted information may include data providing ameasure of the structural integrity of the element. The outputtedinformation may comprise data providing an assessment of the structuralchange in, or damage to, the aircraft component. The information may bein the form of a prediction, for example of the likeliness of a crackforming. The information could for example be in the form of aprediction of the mean time left before a crack will occur. Theoutputted information may include an indication of the expected useablelife-span of the metal element, for example by providing indications ofthe mean time that the element could feasibly remain in service. Theinformation could include indications relating to the structuralintegrity of the element compared to those of a notional averageelement.

The outputted information may include information concerning the sourceor sources of the relevant acoustic emissions. For example, the outputmay include information concerning the location(s) in the element of thesource or sources.

The information outputted is preferably in electronic form, for examplein the form of electronic data.

A multiplicity of acoustic emission sensors may be provided in order toeffect the step of converting acoustic emissions generated in theaircraft component into electronic signals. For example, themultiplicity of acoustic emission sensors may be attached to the metalelement. Alternatively, the multiplicity of acoustic emission sensorsmay be in the form of remote sensors that are able to detect acousticemissions without needing contact with the metal element. For example,the acoustic emission sensors may be in the form of remote laser sourceand detector arrangements, laser light being directed onto the surfaceof the metal element and reflected back to one or more detectors.Preferably, a sufficient number of sensors are provided to enable thelocation of the source of a relevant acoustic emission to be ascertainedby triangulation techniques.

The method may be performed such that at least two of the acousticemission sensors have a fundamental resonant frequency at a firstfrequency and at least two acoustic emission sensors have a fundamentalresonant frequency at a second frequency, the first and secondfrequencies being different. Acoustic emissions at different frequenciesmay thereby be detected. Also, the different acoustic emission sensorsmay be able to provide a frequency-amplitude profile of an acousticemission. The frequency-amplitude profile of an acoustic emission mayfor example be used to assist detection of relevant acoustic emissions.

The metal element may be in the form of a safety critical load-bearingelement that in use is required to be free from cracks. The metalelement need not be made exclusively from metals. For example, the metalelement may be in the form of a metal alloy or mixture containingnon-metallic materials, for example composite material additives. Themetal element may be in the form of any homogeneous structure which isprone to the formation of cracks under fatigue loading, where before theformation of a crack occurs there are structural changes that causeacoustic emissions to be made. The aircraft component may be in the formof any safe-life component on an aircraft. For example the component maybe a component on an aircraft landing gear. The component may be alanding gear leg. The aircraft component may be an engine pylon. Theaircraft component may be a landing gear rib. The aircraft component maybe an aircraft bulk head. The aircraft component will typically have anaverage temperature, during performance of the method, that issubstantially the same as ambient temperature. The average temperatureof the component may for example be less than 100° C.

The method described herein of measuring the structural integrity of anelement may be in the form of a method of measuring the plasticity ofsuch an element. Thus, the information outputted may be a measure of theplasticity of the load bearing metal element.

The methods of the invention described above include both acquisitionand processing of data. The processing of the data may be performed inreal-time soon after the data is acquired. Alternatively, the data maybe stored for subsequent processing at a significantly later time, forexample, during routine periodic maintenance of the metal element. Thus,the present invention further provides a method of acquiring data forsubsequent processing, the data concerning the structural properties ofa safe-life aircraft component comprising a load-bearing metal element,the method including the steps of converting acoustic emissionsgenerated in the metal element into electronic signals, the acousticemissions converted including “relevant acoustic emissions”, namelyacoustic emissions resulting from changes in the structure of theelement, and storing the electronic signals as measurement data in adata store. The data so acquired may then subsequently be processed.Thus there is further provided a method of measuring the structuralintegrity of a safe-life aircraft component comprising load-bearingmetal element, the method including the steps of acquiring measurementdata, and then subsequently processing the data so acquired with aprocessing unit in conjunction with stored reference data that allows ameasure of the structural integrity to be made from the acquiredmeasurement data, and outputting information providing a measure of thestructural integrity of the aircraft component. The measurement data soacquired may include data concerning relevant acoustic emissions, forexample data produced by means of performance of the data acquisitionmethod described immediately above.

The method according to any aspect of the invention described herein isadvantageously performed a multiplicity of successive times in respectof a given metal element of an aircraft component. Preferably, themethod is performed such that the structural integrity of the aircraftcomponent is effectively continuously measured and monitored.

The present invention further provides a processing unit programmed toperform the steps performed by the processing unit of the methodaccording to any aspect of the invention described herein.

The present invention also provides computer software, for example inthe form of a computer software product, that is configured to programmea processing unit to perform the steps performed by the processing unitof the method according to any aspect of the invention described herein.

The present invention further provides computer data, for example in theform of a computer data product, containing reference data for use asthe stored reference data as required by the method according to anyaspect of the invention described herein.

There is also provided a kit of parts including a processing unit and amultiplicity of acoustic emission sensors, the kit of parts being ableto be configured to implement the method according to any aspect of theinvention described herein. The kit of parts may further includecomputer data as described above.

The present invention yet further provides a kit of parts including amultiplicity of acoustic emission sensors and a data storage means forthe storage of measurement data for subsequent processing, the kits ofparts being able to be configured to implement the method according toany aspect of the invention described herein, in which data is storedfor subsequent processing.

In accordance with the present invention there is also provided anapparatus for performing the method of the invention. The apparatusadvantageously includes a safe-life aircraft component comprising aload-bearing metal element, a processing unit, a multiplicity ofacoustic emission sensors, and a reference data store. The apparatus isadvantageously so arranged that

the multiplicity of acoustic emission sensors is arranged to convertacoustic emissions generated in the metal element into electronicsignals,

the processing unit is arranged to receive electronic signals derivedfrom the signals sent by the acoustic emission sensors,

the reference data store includes stored reference data that allows ameasure of the structural integrity of the aircraft component to be madefrom the signals sent to the processing unit,

the processing unit is arranged to process over time the receivedelectronic signals in conjunction with the stored reference data and tooutput information providing a measure of the structural integrity ofthe aircraft component.

The present invention also provides an apparatus for performing anyaspect of the method of the invention described herein of acquiring datafor subsequent processing, the apparatus including a safe-life aircraftcomponent comprising a load-bearing metal element, a multiplicity ofacoustic emission sensors, and a measurement data store. The apparatusis advantageously so arranged that

the multiplicity of acoustic emission sensors is arranged to convertacoustic emissions generated in the metal element into electronicsignals, and

the measurement data store is arranged to receive data signals derivedfrom the electronic signals from the acoustic emission sensors and tostore those signals as measurement data in the measurement data store.

The term structural integrity is used herein in relation to thestructure of the metal element of the aircraft component on the samescale as the size of cracks which are deemed to be too large for safetyreasons on a safe-life structure. The integrity of a structure may bedefined by a measure of the likelihood of the structure containing acrack. The structural integrity of a structure may be defined by ameasure of the likelihood of a crack being formed in the structure, forexample after certain criteria have been satisfied (such as a certaintime having elapsed and/or certain loading conditions having beensatisfied). The structural integrity of a structure may be defined by ameasure of the amount of changes in the submicrostructure thatcontribute to the formation of cracks. The term crack as used herein isintended to cover cracks on a microscopic scale, that is cracks thatcan, once the relevant cross-section is made visible, be identified withthe aid of a microscope. The term “crack” as used herein is alsointended to cover cracks that are able to be readily detectable via theuse of standard non-destructive detection techniques, such as eddycurrent testing. Such techniques reliably enable detection of crackshaving a length of over 0.5 mm. It will of course be understood that thepresent invention advantageously enables detection of cracks and crackformation where the size of the crack is such that the crack would beundetectable using current non-destructive testing techniques. Moreoverthe invention advantageously enables measurements of the structuralintegrity of an aircraft component to be made before cracks form. Itwill also be understood that micro-cracks, in particular micro-crackingon the surface of a metal load bearing element, is not necessarilyeither a crack that significantly prejudices safety or one that needs tobe detected by means of the present invention. However, it should benoted that the present invention may facilitate the detection ofsubmicrostructure changes that contribute to the eventual formation ofcracks and that therefore affect the structural integrity of a safe-lifeaircraft component.

BRIEF DESCRIPTION OF THE DRAWINGS

By way of example, an embodiment of the present invention will now bedescribed with reference to the accompanying drawings, of which:

FIG. 1 shows in plan view a test element with acoustic sensors attachedthereto,

FIG. 2 is a schematic diagram showing a test acoustic emission beingcaused on the test element,

FIG. 3 is a graph showing the S-N curves for the material from which thetest element is made,

FIG. 4 is a graph showing the cumulative number of acoustic emissionsdetected in the test element over time,

FIG. 5 is a graph showing the number of acoustic emissions detected inthe test element over time for 15 consecutive runs, and

FIG. 6 is a schematic diagram showing apparatus for assessing thestructural integrity of a landing gear leg according to the embodiment.

DETAILED DESCRIPTION OF EMBODIMENTS

Experiments were carried out to test the feasibility of the embodimentof the present invention. As the embodiment of the present invention isbased closely on the experiments that were carried out, one of thoseexperiments will now be described in further detail.

The aim of the experiment was to evaluate the effectiveness of anacoustic emission measurement system in detecting early damage in a testelement. Specimen test elements 100 were tested in fatigue, a schematicdiagram of which is shown in FIG. 1. The specimens 100 were made ofhardened 300M steel, and included a hole 102 in their middle. Thespecimens were 350 mm (in the x-direction) by 70 mm (in the y-direction)by 6 mm with a hole 4.4 mm in diameter. The width (in the y-direction)in the waisted middle of the specimens was 26 mm. The specimens wereinstrumented with five sensors S1-S5. The sensor coordinates are givenin Table 1 below:

TABLE 1 Sensor coordinates S1 S2 S3 S4 S5 x = 80, x = 130, x = 230, x =245, x = 270, y = 35 y = 35 y = 30 y = 45 y = 35

The role of sensors S1 and S5 was to enable the filtering out of wavesoriginating from the clamping device used to hold the specimen 100.Sensors S2, S3 and S4 monitor the acoustic emissions coming from thehole and its vicinity. The hole was introduced to ensure the location ofdamage as well as to reduce the scatter factor of the failure of thespecimens.

The resonant frequency of the sensors and their pre-amplifiers was 600KHz. The sensors were surface bonded to the specimens with a noncorrosive silicon rubber. Each pre-amplifier provided an amplificationof 40 dB and a narrow band filtering of the sensor output. The acousticemission measurement system was connected to the pre-amplifiers viacoaxial cables. The load applied was read by the acoustic emissionmeasurement system using one of its non acoustic data inputs via acoaxial cable.

Before fatiguing of the specimen was commenced the sensors S1 to S5 werecalibrated. This operation consisted of verifying that each sensor wasin good functioning order and that the sensor was properly bonded to thespecimen. The calibration procedure also included a step of makingacoustic emission measurements and evaluating the group velocitythroughout the specimen. These group velocity measurements were usedlater to ascertain the location of the acoustic emission sources. Pencillead was broken (Hsu-Nielson source) on the specimen surface, theacoustic emission measurement system measuring the time difference offlight, ΔT_(i), namely the travel time difference from the first hitsensor to the i_(th) hit sensor. With reference to FIG. 2, if theseparation between the acoustic emission AE and the first sensor S1 isdistance a₁ and the separation between the acoustic emission AE and thesecond sensor S2 is distance a₂, then the velocity is calculated asb/ΔT, where b=a₂−a₁, and ΔT is the time between the acoustic emissionbeing detected at sensor S1 and the acoustic emission being detected atsensor S2.

The measured velocity was approximately 5 km/s, which is in agreementwith the theoretical value given by the dispersion curves. The specimen1 was then subjected to loading in the form of a sinusoidal cycle ofconstant amplitude 13 Hz frequency. FIG. 3 shows the S-N curves for the300M material used to determine the maximum stress level to apply on thespecimen to reach a specific load cycle for a given stress ratio. Theloading was applied in a series of 15 loading runs, the loading profileapplied to a particular specimen being as shown in Table 2, set outbelow:

TABLE 2 Details of profile of loading applied to a specimen CumulativeLoading Peak Min. No. of cycles at Run Load Load cycles end of run 131.5KN 3.1KN 335,960 335,960 2 31.5KN 3.1KN 166,670 502,630 3 31.5KN3.1KN 272,265 774,895 4 31.5KN 3.1KN 432,005 1,206,900 5 37.4KN 3.1KN191,050 1,397,950 6 37.4KN 3.7KN 146,650 1,544,600 7 37.4KN 3.4KN150,010 1,694,610 8  42KN 4.3KN 220,000 1,914,610 9  42KN 4.3KN 230,0002,144,610 10   42KN 4.3KN 209,990 2,354,600 11   42KN 4.3KN 90,0102,444,610 12   42KN 4.3KN 220,160 2,664,770 13   42KN 4.3KN 190,0902,854,860 14a 52.5KN 5.2KN 59,740 2,914,600 14b 21.2KN 2.1KN 80,9502,995,550 15a  26KN 2.6KN 39,050 3,034,600 15b 31.5KN 3.1KN 38,4103,073,010

The acoustic emission measurement system was able to measure the loadapplied in the fatigue test by means of non-acoustic measurements.During the experiment, the specimen was inspected for cracks after eachrun using both a microscope and NDT (Eddy Current) techniques. Acousticemissions detected by the acoustic emission measurement system overthreshold amplitude were measured and counted.

FIG. 4 illustrates a graph showing the cumulate burst count on they-axis against time (that is time during loading, the time between runsbeing ignored) on the x-axis. The x-axis of the graph of FIG. 4 isdivided into 15 segments, each segment representing a load run, so thatthe single curve on the graph represents the cumulative burst count andtime passing as measured from the start of the first load run (run 1).FIG. 5 also shows the cumulative burst count against time for the samedata as that represented by FIG. 4, but in FIG. 5, there are 15 separatecumulative burst curves, one for each loading run, the curves showingthe cumulative burst count and time passed as measured from the start ofthe load run for that curve.

As can be seen, the specimen was loaded at various load levels, from31.5 KN to 52.5 KN. A crack of length of 1.5 mm was observed as havingbeen initiated during load run 14, that is after about 2.9 millioncycles.

During the earlier runs, the gradient of the curves of FIGS. 4 and 5 isabout 0.1-0.2 burst/s. The specimen generated a gradient about 0.1burst/s between 0.3 million cycles and 1.43 million cycles correspondingto zones 2 to 5. The gradient increased to 0.2 burst/s between 1.43million cycles and 2.13 million cycles corresponding to zones 8 to 10.The gradient further increased to 0.3 burst/s between 2.4 million cyclesand 2.8 million cycles corresponding to zones 11 to 13. A significantshift in gradient (0.6 burst/s) was noticed between 2.86 million cyclesand 2.93 million cycles corresponding to the load run (zone 14) wherethe crack was detected by a microscope and also by using NDT (eddycurrent) techniques. It will also be noted that the gradient dropped toalmost zero immediately before the failure (zone 15) of the specimen.

As a result of the experiments that have been conducted, a method ofmonitoring the structural integrity of landing gear for an aircraft hasbeen proposed. The method and the apparatus for implementing thisproposal will now be described with reference to FIG. 6, which shows ablock diagram illustrating the function of the proposed embodiment.

FIG. 6 shows a landing gear leg 110, in which there are embedded variousacoustic emission sensors, of which only four are shown, S1-S4. Outputsfrom these sensors are fed via analogue to digital converters (notshown) to an acoustic emission measuring system 112. The signals fromthe sensors S1-S4 are received at a comparator/filter system 114, whichassesses whether the magnitude and frequency of the acoustic emissionsreceived from the sensors are within preset criteria so as to be deemedas acoustic emissions (hereinafter “significant acoustic emissions”)resulting from changes within the microscopic structure of the landinggear 110, as opposed to acoustic emissions resulting from other sources.The parameters defining any significant acoustic emission are thenextracted for use in analyzing the structural integrity of the metalload bearing structure of the landing gear leg. The sensors andelectronic equipment used to detect and analyse acoustic emissions arewell known in relation to monitoring of cracks in metals and suchapparatus may be used to implement the present embodiment. One suchapparatus is described in WO 01/94934, the contents of which (inparticular the contents concerning the apparatus and methods used todetect and analyse acoustic emissions as described in that document withreference to the drawings of that document) are incorporated herein byreference thereto.

The apparatus of the invention is used to make an assessment of thestructural integrity of the landing gear leg, over time, by means ofvarious methods of analysis of the measure of cumulative bursts overtime. In use, a processor 116 of the measurement system 112 receivesdata from the comparator/filter 114 concerning extracted parametersdefining the acoustic emissions judged by the comparator/filter 114 asbeing significant acoustic emissions. This data is then analysed inconsideration of data stored in a memory store 120 that allows theprocessor 116 to effectively compare the real-time data with data storedin the memory 120 so that statistically valid conclusions can be drawnconcerning the structural integrity of the landing gear 110. Theprocessed data and results are stored in a further memory store 118 fordownloading during routine maintenance of the aircraft.

The method can be considered as plotting graphs similar to that shown inFIGS. 4 and 5 and analyzing various characteristics of such graphs. Ashas been established by experiment, there appear to be many ways inwhich the conditions that facilitate crack formation can be correlatedto data extracted from measuring significant acoustic emissions.Indications of the structural integrity are provided by means ofcomparing the acoustic emissions data retrieved in use concerning thelanding gear with a variety of thresholds and criteria that have beenpre-set by means of prior experimentation and/or mathematical modelling.The criteria against which the structural integrity of the landing gearis compared, in this embodiment, consist of monitoring the following:

-   -   the absolute cumulative burst,    -   the number of bursts over a range of different time periods,    -   the burst rate,    -   the integral of cumulative bursts over time,    -   the above parameters when weighted by burst peak amplitude (so        that more energetic acoustic emissions are given more weight        than less energetic emissions), and    -   the above parameters when considered over time when grouped by        the activity of the aircraft.

In each case, the data analysed is compared against the stored referencedata and a result is issued with an associated statistical probability.For example, the result might be in the form that the data recordedindicates that 1% of landing gears having the same data would be beyond75% of the working life of the gear, and that 0.1% of landing gearhaving the same data would be beyond 80% of the working life of thelanding gear. The result might also be in a form that states that 1% oflanding gear having the same data would be beyond 23% of the expectedtime till first crack is detected, and that 0.1% of landing gear havingthe same data would be beyond 28% of the expected time till first crackis detected. During maintenance of the aircraft such results may be usedto decide when a particular landing gear leg should be replaced, withthe benefit of increased confidence in the structural integrity of alanding gear and possibly the benefit of enabling landing gear to be inservice for longer than is now safely possible.

The criteria for assessing the structural integrity of the landing gearleg 110 will now be briefly discussed in turn with reference to thegraphs shown in FIGS. 4 and 5. Whilst various thresholds and numbers arediscussed with reference to FIGS. 4 and 5, it will of course beappreciated that FIGS. 4 and 5 correspond to data relating to a specimentest element.

Absolute Cumulative Burst

The cumulative burst count until a crack appears is similar foridentical specimens. Thus, a threshold cumulative burst count can beset, over which threshold the landing gear leg should be replaced.

Number of Bursts Over a Range of Different Time Periods and Burst Rate

As can be seen from FIGS. 4 and 5, the gradient of the curve generallyincreases as the curve gets closer to the instant at which a crack firstappears. Thus, a threshold burst rate can be set, over which thresholdthe landing gear leg should be replaced. It will however be appreciatedthat the rate can increase to a level comparable to that reachedimmediately before a crack appears even though the material is not veryclose to a state in which cracks might appear. For example, consider thegradient of the curves corresponding to test runs 7 and 14. As can beseen more clearly in FIG. 5, the gradient of the latter part of thecurve corresponding to test run 7 is almost as steep as the gradient ofthe curve corresponding to test run 14, even though a crack firstappeared during test run 14, and test run 7 might be seen ascorresponding to 50% of the maximum possible useable lifetime of thespecimen. Thus, advantageously, other criteria are used to reduce thechance of a landing gear leg being withdrawn from service prematurely.Such criteria can include assessing the absolute cumulative burst countin conjunction with the gradient. For example, gradients of the curve ata given level but corresponding to a cumulative burst count being belowa threshold cumulative burst count may effectively be ignored, whereasgradients of the curve at the same given level but corresponding to acumulative burst count above the cumulative burst count threshold may beconsidered as warranting replacement of the landing gear leg.

The average gradient over a pre-selected interval can also be monitored.It will be seen that curve 14 has a steep gradient that is sustainedover a number of bursts over 7000, whereas the steep section of curve 7last for only about 4000 bursts. Thus, there may be set a thresholdaverage gradient (for example in this case being 0.625 bursts/second)which must be exceeded when measured over a certain number of bursts,for example in this case, 5000 bursts. Line 124 is a line that spans5000 bursts and which has a gradient of 0.625 bursts/second, whereasline 126 is a line that spans 5000 bursts has a gradient steeper than0.625 bursts/second. Alternatively, the average gradient may be requiredto be maintained for a given length of time. For example, an averagegradient of greater than 0.93 bursts/second may need to be maintainedfor at least 2250 seconds. Line 120 is a line that spans 2250 secondsand has a gradient of 0.93 bursts/second, whereas line 122 is a linethat spans 2250 seconds and has a gradient steeper than 0.93bursts/second.

It will of course be appreciated that a number of such criteria can becombined such that the effective test is whether over any given timeinterval the curve of the number of bursts against time crosses apre-set boundary. Such a notional boundary is illustrated by the shadedarea 128 in FIG. 5. Thus, periodically (at time T, say) the boundarycriteria are applied to the curve as would be drawn for the periodT−T_(test) until T, where T_(test) is a constant time period of, say,2000 seconds (the origin of the graph being at the point wherex=T−T_(test) and y=0). If any part of the curve crosses the boundaryinto the shaded area 128 the processor will decide that the landing gearleg needs replacing.

Integral of Cumulative Bursts Over Time

Again, the area under the curve of cumulative burst count over time canbe monitored and can provide indications of the integrity of thestructure of the landing gear leg. The integral over lifetime can bemonitored as can the integral over shorter periods of time.

Weighting of Measurements

All of the above methods of monitoring the structural integrity of thelanding gear leg rely on counting significant acoustic emissions,irrespective of their location, amplitude, duration or otherparameters/characteristics of the acoustic emission. More sophisticatedcalculations can be made to weight the burst count total in view of oneor more parameters or characteristics of the acoustic emission detected.For example, each burst measured could be weighted by the burst peakamplitude. Thus higher energy acoustic emissions (corresponding tocomparatively greater change to the internal structure of the landinggear leg) are generally given greater weight than lower energy acousticemissions. Such weighting could effectively replace, in part at least,the filtering and selecting step performed by the comparator/filter 14,in that acoustic emissions that would previously be discounted as notqualifying as a significant acoustic emission are now accounted but areweighted to have less effect on the analysis carried out.

Separation of Activity of the Aircraft

Because the loading of the landing gear differs significantly accordingto the activity of the aircraft, the measurements made can either beweighted according to activity or measurements could be made andanalysed in groups according to the activity of the aircraft. Forexample, the landing gear is subjected to loads during taxiing, landingand takeoff. At other times the loading on the landing gear is notsignificant enough to warrant continued monitoring of acousticemissions. Giving that loading during landing is greater than duringtaxiing, the acoustic emissions detected during landing can be givengreater weight than during taxiing. Alternatively, separate logs can bemade, and/or different rates of data acquisition may be used, formeasurements made during taxiing, landing and takeoff, respectively.

Combination of Methods

A variety of different methods for analyzing the acoustic emissionsdetected are described above and it will be appreciated that acombination of a plurality of such methods may be implemented. Thechoice of the methods implemented will depend on various factorsincluding the reliability and in particular the statistical validity ofthe methods actually employed. Such choices can be determined andverified by means of routine experimentation and testing.

In the embodiment described above, the data processed by thecomparator/filter 114 is passed to a processor 116 for real-timeprocessing, the results of which being stored in a memory 118. Oneadvantage of such a system is that if, for whatever reason, theprocessor determines that the structural integrity of the landing gearleg rapidly deteriorates a warning can be made immediately. However, itis also acceptable for the data from the comparator/filter 114 to beprocessed separately. For example, the measurement system 112 could beprovided without the processor 116 and the memory store 120 ofpreviously recorded data for comparison with the measured data. In sucha case, the data from the comparator/filter 114 would be simply storedin memory 118 for downloading during maintenance of the aircraft, suchthat the processing and analysis of the data is performed separatelyfrom the aircraft. Such a proposal would reduce complexity of theaircraft processing systems and would also have the capacity to reduceweight slightly.

Whilst the present invention has been described and illustrated withreference to a particular embodiment, it will be appreciated by those ofordinary skill in the art that the invention lends itself to manydifferent variations not specifically illustrated herein. For thatreason, reference should be made to the claims for determining the truescope of the present invention. By way of example, certain furthervariations to the above-described embodiment will now be described.

The peak amplitudes of the acoustic emissions at the source of eachacoustic emission may be monitored. The peak amplitude at the source maybe calculated by means of ascertaining the location of the source ofeach acoustic emission. The location of the source may be ascertained bymeans of triangulation. The general trend of the peak amplitudes ofacoustic emissions may be monitored over time. For example, it isthought that the peak amplitudes of acoustic emissions may in certainapplications first follow a general upward trend and thereafter decreasefollowing a general downward trend, after which crack initiation occurs.Thus, in an alternative embodiment of the invention, a prediction ofimminent crack formation is made when the trend in peak amplitudes ofacoustic emissions (the calculated peak amplitude of the acousticemissions at their respective sources) exceeds a first preset thresholdand then subsequently decreases below a second preset threshold. Such amethod may in itself be sufficient to make a reasonably accurateprediction of crack initiation.

As an alternative, prediction of crack initiation may be based primarilyon monitoring the rate of relevant acoustic emissions. For example, themethod may include monitoring the general trend of the change in therate of relevant acoustic emissions. Thus, in this alternativeembodiment of the invention, a prediction of imminent crack initiationis made when the trend in rate of acoustic emissions exceeds a firstpreset threshold and then subsequently decreases below a second presetthreshold.

A further variation comprises monitoring both the rate of acousticemissions and the cumulative number of acoustic emissions. Once bothmonitored parameters exceed preset thresholds (or meet other presetcriteria) the aircraft component being monitored may be deemed to be inneed of urgent replacement.

Where in the foregoing description, integers or elements are mentionedwhich have known, obvious or foreseeable equivalents, then suchequivalents are herein incorporated as if individually set forth.Reference should be made to the claims for determining the true scope ofthe present invention, which should be construed so as to encompass anysuch equivalents. It will also be appreciated by the reader thatintegers or features of the invention that are described as preferable,advantageous, convenient or the like are optional and do not limit thescope of the independent claims.

The invention claimed is:
 1. A method of measuring the structuralintegrity of a safe-life aircraft component on an aircraft, thecomponent comprising a load-bearing metal element that is free fromcracks, the method including the steps of: converting acoustic emissionsgenerated in the metal element into electronic signals, the acousticemissions converted including relevant acoustic emissions resulting fromchanges in the structure of the element that make the element moresusceptible to the formation of cracks, sending the electronic signalsto a processing unit, and with the processing unit processing over timethe signals in conjunction with stored reference data that allows ameasure of the structural integrity to be made from the signals sent tothe processing unit, and outputting information providing a measure ofthe structural integrity of the aircraft component before any cracksoccur.
 2. A method according to claim 1, wherein the informationoutputted comprises a prediction concerning the likely occurrence ofcrack initiation.
 3. A method according to claim 2, wherein theprediction comprises an indication of the predicted mean time leftbefore a crack will occur.
 4. A method according to claim 1, wherein thestep of processing the signals includes calculating the cumulativenumber of relevant acoustic emissions over a period.
 5. A methodaccording to claim 4, wherein the step of processing the signalsincludes a step of effectively comparing the cumulative number ofrelevant acoustic emissions with a pre-set threshold.
 6. A methodaccording to claim 1, wherein the step of processing the signalsincludes a step of monitoring the number of relevant acoustic emissionsover a pre-set period.
 7. A method according to claim 1, wherein thestep of processing the signals includes a step of calculating acharacteristic relating to the rate of relevant acoustic emissions.
 8. Amethod according to claim 1, wherein the step of processing the signalsincludes taking into account an indication of the proximity of thesources of the respective relevant acoustic emissions relative to eachother.
 9. A method according to claim 1, wherein the informationproviding a measure of the structural integrity of the aircraftcomponent comprises an indication whether or not the aircraft componentis deemed to be in need of further action.
 10. A method according toclaim 9, wherein the further action is further non-destructive testingto obtain an assessment of the structural integrity of the aircraftcomponent, such testing not being dependent on the monitoring ofacoustic emissions.
 11. A method according to claim 10, wherein themethod includes implementing the further action.
 12. A method accordingto claim 9, wherein the further action comprises replacing at least apart of the aircraft component.
 13. A method according to claim 12,wherein the method includes implementing the further action.
 14. Amethod according to claim 9, further comprising implementing the furtheraction.
 15. A method according to claim 14, wherein the further actionis further non-destructive testing to obtain an assessment of thestructural integrity of the aircraft component, such testing not beingdependent on the monitoring of acoustic emissions.
 16. A methodaccording to claim 1 including providing a multiplicity of acousticemission sensors for effecting the step of converting acoustic emissionsgenerated in the aircraft component into electronic signals.
 17. Amethod according to claim 1, wherein the aircraft component is in theform of a component on an aircraft landing gear.
 18. A processing unitprogrammed to perform the steps performed by the processing unitaccording to claim
 1. 19. A non-transitory computer readable mediumcontaining therein software that is configured to program a processingunit to perform the steps performed by the processing unit in the methodaccording to claim
 1. 20. A non-transitory computer reading mediumcontaining therein reference data for use as the stored reference dataas required by the method according to claim
 1. 21. A kit of parts,comprising a processing unit and a multiplicity of acoustic emissionsensors, the kit of parts being configured to implement the methodaccording to claim
 1. 22. A kit of parts according to claim 21, furthercomprising a computer reference data for use as the stored referencedata as required by the method.
 23. A method according to claim 1,wherein the information providing a measure of the structural integrityof the aircraft component before any cracks occur comprises anindication that the aircraft component is in need of further action; andthe method further comprises implementing the further action in responseto said indication.
 24. A method according to claim 23, wherein thefurther action is at least one selected from the group consisting ofnon-destructive testing to obtain an assessment of the structuralintegrity of the aircraft component, such testing not being dependent onthe monitoring of acoustic emissions; testing the load-bearing metalelement; performing a routine test on the load-bearing metal elementmore often; replacing the aircraft component or the load-bearing metalelement; and repairing the aircraft component or the load-bearing metalelement.
 25. A method of acquiring data for subsequent processing, thedata concerning the structural properties of a safe-life aircraftcomponent on an aircraft, the component comprising a load-bearing metalelement that is free from cracks, the method comprising: convertingacoustic emissions generated in the metal element into electronicsignals, the acoustic emissions converted including relevant acousticemissions resulting from changes in the structure of the element thatmake the element more susceptible to the formation of cracks, assessingwhich of the electronic signals correspond to relevant acousticemissions, and storing the electronic signals that have been assessed ascorresponding to relevant acoustic emissions as measurement data in adata storage.
 26. A kit of parts, comprising a multiplicity of acousticemission sensors and a data storage unit for the storage of measurementdata for subsequent processing, the kits of parts being configured toimplement the method according to claim
 25. 27. A method of measuringthe structural integrity of a safe-life aircraft component on anaircraft, the component comprising a load-bearing metal element, themethod comprising the steps of: arranging at least one acoustic emissionsensor on or around the metal element that is free from cracks,acquiring, by the at least one acoustic emission sensor, measurementdata, including data concerning relevant acoustic emissions generated inthe metal element as a result of changes in the submicrostructure of themetal element that contribute to the formation of cracks and make theelement more susceptible to the formation of cracks, processing the dataso acquired with a processing unit in conjunction with stored referencedata that allows a measure of the structural integrity of the aircraftcomponent to be made from the acquired measurement data, and outputtinginformation providing a measure of the structural integrity of theaircraft component before any cracks occur.
 28. A processing unitprogrammed to perform the steps performed by the processing unitaccording to claim
 27. 29. A non-transitory computer readable mediumcontaining therein software that is configured to program a processingunit to perform the steps performed by the processing unit in the methodaccording to claim
 27. 30. An aircraft, comprising: a crack-freesafe-life aircraft component forming a load-bearing metal element of theaircraft, and an apparatus for performing the method of claim 1 on thesafe-life component, the apparatus comprising a processing unit, amultiplicity of acoustic emission sensors, and a reference data storage,wherein the multiplicity of acoustic emission sensors is arranged toconvert acoustic emissions generated in the metal element intoelectronic signals, the processing unit is arranged to receiveelectronic signals derived from the signals sent by the acousticemission sensors, the reference data storage includes stored referencedata that allows a measure of the structural integrity of the aircraftcomponent to be made from the signals sent to the processing unit, andthe processing unit is arranged to process over time the receivedelectronic signals in conjunction with the stored reference data and tooutput information providing a measure of the structural integrity ofthe aircraft component.
 31. An aircraft, comprising: a crack-freesafe-life aircraft component forming a load-bearing metal element of theaircraft, and an apparatus for performing the method of claim 25 on thesafe-life component, the apparatus comprising a multiplicity of acousticemission sensors, and a measurement data storage, wherein themultiplicity of acoustic emission sensors is arranged to convertacoustic emissions generated in the metal element into electronicsignals, and the measurement data storage is arranged to receive datasignals derived from the electronic signals from the acoustic emissionsensors and to store those signals as measurement data in themeasurement data storage.